This invention relates generally to gas turbine engine cooling circuits and more particularly, to methods and apparatus for cooling gas turbine engine nozzle assemblies.
Gas turbine engines include combustors, which ignite fuel-air mixtures, which are then channeled through a turbine nozzle assembly toward a turbine. At least some known turbine nozzle assemblies include a plurality of nozzles arranged circumferentially within the engine downstream from the combustors. Each nozzle includes a hollow airfoil vane that extends between integrally-formed inner and outer band platforms. The nozzles are cooled by a combination of internal convective cooling and gas side film cooling.
Each hollow airfoil is supplied cooling air through an internally-defined cavity that is bounded by a pair of connected sidewalls. Cooling of engine components, such as components of the high pressure turbine, is necessary due to thermal stress limitations of materials used in construction of such components. Typically, cooling air is extracted air from an outlet of the compressor and the cooling air is used to cool, for example, turbine nozzles. At least some known turbine nozzles include cooling circuits within the cavity which define flow paths for channeling cooling air flow through the cavity for cooling the airfoil prior to the air flow being discharged downstream through trailing edge slots defined within the airfoil.
Within at least some known airfoil cavities, a serpentine shaped path or channel having multiple chamber passes is defined. Within channel flow circuits, the heat transfer coefficient of coolant flowing through a channel is a function of the local flow velocity in the circuit. Because the metal temperature distribution of a typical vane airfoil is such that the trailing edge is significantly hotter than a temperature of the bulk of the airfoil, at least some known airfoils use turbulence promoters such as pins, turbulators, and other roughening devices to increase the heat transfer coefficient of the coolant flowing through the channel. However, within channel flow circuits, as a portion of the airflow is channeled aftward through the trailing edge slots, a region of low heat transfer coefficient may form near the end of the cooling path. To facilitate cooling in such regions, at least some known airfoils use local film cooling. However, in advanced applications, in which the engine may be operated under extreme heat load conditions, a region of low cooling at the end of a cooling circuit may not be solved by local film cooling, and as a result, may limit the operating range and acceptable applications of the engine.
In one aspect of the invention, a method for fabricating a nozzle for a gas turbine engine is provided. The nozzle includes an airfoil, and the method includes forming the airfoil to include a suction side and a pressure side connected at a leading edge and a trailing edge such that a cooling cavity and a cooling circuit are defined within the airfoil, wherein the suction side and the pressure side extend radially between a tip and a root. The method also includes forming a plurality of cooling slots within the airfoil that extend from the cooling circuit towards the airfoil trailing edge, and forming a control vane within the cooling circuit to facilitate maintaining a substantially constant cooling effectiveness within the cooling circuit.
In another aspect, a turbine nozzle for a gas turbine engine is provided. The nozzle includes an airfoil vane that includes a first wall, a second wall, a plurality of trailing edge cooling slots, and a cooling circuit that extends between the first and second walls. The cooling circuit is upstream from the trailing edge cooling slots for channeling cooling air to the trailing edge cooling slots. The cooling circuit includes at least one control vane extending between the first and second walls. The control vane is arcuate and extends upstream from the trailing edge cooling slots for maintaining a substantially constant cooling effectiveness within said cooling circuit.
In a further aspect of the invention, an airfoil for a gas turbine engine nozzle is provided. The airfoil includes a root, a tip, a cooling circuit, a plurality of trailing edge cooling slots, and a convex sidewall and a concave sidewall that are connected at a trailing edge. The plurality of trailing edge cooling slots extend from the cooling circuit towards the airfoil trailing edge. Each sidewall extends between the root and tip. The cooling circuit is defined between the sidewalls and includes a plurality of pins and a control vane. The plurality of pins and the control vane extend between the sidewalls and define a flowpath for channeling cooling air through the cooling circuit into the trailing edge cooling slots. The control vane is configured to facilitate maintaining a substantially constant cooling effectiveness within the cooling circuit.